Contrarotating turbojet engine having independent bearing supports for each turbocompressor



Sept. 2, 1952 E, HUNSAKER 2,608,821

CONTRAROTATING TURBOJET ENGINE HAVING INDEPENDENT BERNG SUPRORTS FOREACH TURBOCOMPRESSOR Filed OC. 8, 1949 Patented Sept. 2, 1952 UNITED fSTATE v 2,608,821,l Y

coNTnAnoTArING TUn'BolJE'I nncmirf` 1. Y HAVING INDEPENDENT BEARING surroars ron EACH TUaBocoMrREsson vEugene L. Hunsaker, Manhattan Beach,Calif., assigner, by mesne assignments, to General," Electric Company,Schenectady, N. Y., a corporation of New York Application october s,194e, sentirne. 120,247,

1 Claim. 1

My invention relates to axial flow compressorturbines of the gaseouscombustion type, and more particularly to an improved compressor.-turbine suitable for use in aircraft driven at high subsonic orsupersonic speeds.

When combustionV turbines are utilized in such aircraft, high airowcapacity is necessary to increase the engine thrust coecient and to de,-crease the external drag coelicient. At present, the required airflowsare obtained by the use of high blade speeds in axial new compressors ofcompressor-turbine combinations used to drive the high speed airplanes.

High blade speeds `Vresult in large values of centrifugal stress on the.rotating blades as Stress (centrifugal):

l( blade tip speed)2 X blade length constant coinpressordiarneter Thisstress relationship in the compressor limits the blade length and hencelimits flow area and capacity fora given diameter of compressor. Adirectly coupled turbine Will also have large values of blade tip speed,and hence high values of centrifugal stress. The physicalcharacteristics or" the hot'turbine bladesparticularly limits bladelength, and hence airflow rate ior a. given turbine diameter.

It is an object of the present invention to provide a compressor-turbinecombination which is capable of higher air capacities than areconventionally obtainable with compressors. using high blade speeds. andsupersonic air velocities relative to the blades.Y y

It is another object of thev present invention to providean improvedcompressor-turbine combination having relatively low blade tip speeds.

It is still another objectof the invention to provide acompressor-turbine..combinationjhaving relatively long blade lengths,with aconsequent signiicant increase in airow capacity.

It is a further objectoi thepresent invention to provide acompressor-turbine,- combination of the combustion type suitable fordriving aircraft at high subsonic and supersonic speeds.

A still further object of the invention is to provide an improvedcompressor-turbine in which high compression ratios can be produced in asingle compressor stage, while maintaining high air capacity andrelatively low blade tip speeds.

Another object of the present invention is to provide higher compressoreiciencies in compressor-turbines operating at supersonic relativevelocities.

Briefly, the present invention involves a com- (Cl. (iO-35.6).

bustionturbine 'engineVv -having contra-rotating compressorblade..discs..directly ,driven by contrarotating turbine blade ,discsfThe terms ycompressor rotor blades and turbine rotor blades are used .inthe conventional sense of meaning bladeshaving the.. proper. contoursVfor rotating blades of. compressorsandturbines of compressor-turbineengines;-- whereastheterms compres-` sor.- counter-rotating,blades ndturbine counterrotating` blades are usedixr an unconventional senseindicating only thatthe counter-rotating blades have respective propercontours. .for v compressor stator blades and! turbine stator blades.

even though rotating. i Y

My present.v invention provides for ay compressor-turbine in whichthecompressorjbladesfof, the

stator typev are; contrafrotait-xi with respect to the compresser-rcter`bladesjso.- that.Y the compressor-rotor speed` canbey decreased for: a'given overall stage pressure ratio. I prefer to drive the;contra-rotating compressor blades by a turbine-v disc havingrturbinerotor bl'ading thereon, working in conjunction-with a.contra-rotating disc having turbineystatorftypeblades thereon, this discbeingconnectedfxto drivethe compressor-rotorblading. "The compressorrotor blading is preferably directly connected' to .a turbine. discYhaving turbine.' stator. type bladesthereion, so that both compressor:and'. turbine comprise, contrarotat-ing blading'. .In;this wiay thefrtipspeeds of both compressor anduturbine .bladin'g are relatively low.,and: anv increased. air: .capacity is obtained. forY a. givenv engine'.diameter.

lnzthisfdesigrr, thefstator .type blades are contra-rotated, withrespect' to'thezrotor type blades and the ,bla-1deboundary. layerenergizedV through this relative rotation.: suppresses separation of'thef fiowfrom. thefstator; type; blades and:

improves efficiency.

My'invention. will be;4 morcef: fully understood by reference: to the.drawing-3. showing in` diagrammatic for-1n,Y a longitudinal. sectionalview of a combustion turbo-jet engine embodying a preierred'f form: ofvthe: present. invention.

Referring; to'A the.drawing,.an= outer turbine casing. IV is-providedhaving Yat one end an airflowv inlet 2 and at the opposite end, a jetoutlet V3. Attached to casing I just inside of the airflow inlet, aplurality of radial guide vanes 4 are provided, extending inwardly tosupport an inlet cone 6, which is hollow, and in turn supports aplurality of inlet bearing hangers 1 supporting an outer race 8 of aforward bearing 9.

Inside of the jet outlet 3 a plurality of radial tail cone hangers l0are provided, supporting a tail cone II. Inside of tail cone II aplurality of outlet bearing hangers I2 are provided, supporting theouter race I4 of a rear bearing I5. Journalled in bearings 9 and I5 isan inner cylindrical rotor i6. The forward endof theV inner rotor I6 isprovided with a compressor rotor disc I'I on which are mounted aplurality of compressor rotor blades I8 just rearwardly of guide vanes4. Attached to the rear of inner rotor I6 is a turbine counter-rotatingdisc lI9 provided with blades 2D of turbine stator type, these bladesbeing the rearmost of the turbine assembly.

Between the compressor rotor `blades I8 and the turbine stator typeblades 20, vinside of casing I and adjacent the inner surface thereof,is positioned a, ring of combustion chambers 2| of any conventionaltype, each supplied with fuel through a, fuel inlet jet 22. y

Forwardly of the combustion chamber ring, attached to casing I andextending radially inwardly is a ring of stationary compressor outletguide vanes 2 4 supporting an inner shroud 25, forming an annular spacefor combustion chambers 2I. Y

Rearwardly of the combustion chambers 2l, attached to casing l andextending radially inwardly is a ring of stationary turbine inletnozzles 26 also supporting the Ainner shroud 25.V Extending inwardlyfrom kshroud 25 at the guide.

vane and nozzle stations, are a plurality of radially arranged struts 21and 28 supporting an outer rotor 30 rotating coaxially with inner rotorI6 on front and rear outer rotor bearings 32 and 34 respectively. ,l l YCompressor Y rotor blades I8'T'an'd' stationary guide vanes 24 arespaced to permit insertion therebetween of a plurality of compressorstator type blades 35 attached tothe forward end of outer rotor 3U, bycompressor counter-rotating disc 36.

Similarly at the turbine, nozzles 26 and Yturbine stator type blades 20are spaced to provide for a plurality of turbine rotor blades 37 mountedon a rear rotor disc 38 of youter rotor 30, Thus both the compressor androtor have two sets of rotating blades. As one set of blades of bothcompressor and turbine are shaped as rotor type blades, and sinceanother set of blades in each unit are shaped as stator type blades, theinner and outer rotors will be contra-rotating, with both rotor andstator type'blades of the compressor beingv driven bythe turbine blades,but in opposite directions. i It will be apparent to one skilled in theart that, for a given overall stage pressure ratio, the rotor speed isdecreased by the use of such an arrangement. lThe decreased speedpermits greatercompressor and turbine blade length, thereby increasingair capacity for a given engine diameter.

While in order to comply with the statute, the invention has beendescribed in language more or less specific as to structural features,it is to be understood that the invention is not limited to the specicfeatures shown. but that the means and 'construction herein disclosedcomprise a preferred form oi' putting the invention into eiect, and theinvention is, therefore, claimed in any of its forms or modificationswithin the legitimate and valid scope of the appended claim.

What is claimed is:

A gas turbine combination comprising an outer casing having an airentrance aperture and a jet aperture, a nose cone in said entranceaperture, a plurality of guide vanes extending from said casing andsupporting said nose cone, a forward bearing supported by and withinsaid nose cone, a tail cone in said jet aperture, a plurality of strutsextending from said casing and supporting said tail cone, a rear bearingsupported by and within said tail cone, an inner rotor journalled insaid forward and rear bearings, a first compressor disc on the forwardend of said inner rotor, a plurality of blades of compressor rotor typeon said first compressor disc, a second turbine disc on the rear end ofsaid inner rotor, a plurality of blades of the turbine stator type onsaid second turbine disc, a shroud positioned in said casing betweensaid discs, supports extending from said casing and holding said shroudin said casing to provide an annular space between said shroud andcasing, front and rear shroud bearings inside of and supported by saidshroud, an outer rotor supported mechanically independent of said innerrotor and mounted in said latter shroud bearings to rotateconcentrically with and spaced from said inner rotor, a secondcompressor disc on the forward end of said outer rotor, a plurality ofblades of the compressorstator type on said forward disc, said latterblades being positioned adjacent and to the rear of the blades on saidrst compressor disc, a rst turbine disc on the rear end of said outerrotor, a plurality of blades of turbine rotor type on said rst turbinedisc, said latter blades being positioned adjacent and forward of theblades on said second turbine disc, and combustion chamber means in theannular space between said shroud and said casing.

EUGENE L. HUNSAKER.

REFERENCES CITED The following references are of record in the iile ofthis patent:

UNITED STATES PATENTS Number Name Date 2,360,130 Heppner Oct.` 10, 19442,396,911 Anxionnaz Mar. 19, 1946 2,404,767 Heppner July 23, 19462,405,723 Way Aug. 13, 1946 2,409,176 Allen Oct. 15, 1946 2,430,399Heppner Nov. 4, 1947 2,476,179 Cameron July 12, 1949 2,483,401 Cole Oct.4, 1949 2,505,660 Baumann Apr. 25, 1950 2,563,744 Price Aug. 7, 19512,575,682 Price Nov. 20, 1951 Y FOREIGN PATENTS Number Country Date879,123 France Nov. l0, 1942

